1. Field of the Invention
The present invention relates generally to a turbine blade, and more specifically to low flow cooling demand for the airfoil and the tip regions.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a hot gas flow is developed in the combustor from the burning of a fuel with compressed air from the compressor and then passed through a multiple staged turbine to produce mechanical power. In an aero engine, the mechanical power drives the rotor shaft that is connected to a bypass fan. In an industrial gas turbine engine, the rotor shaft is connected, to an electric generator that will produce electrical power. In both engines, the engine efficiency can be increased by passing a higher temperature gas into the turbine. However, the turbine inlet temperature is limited to the material properties of the first stage turbine airfoils, these airfoils being the stator vanes and the rotor blades.
Complex internal airfoil cooling passages have been proposed to provide high levels of airfoil cooling using a minimal amount of cooling air. Higher turbine inlet temperatures are obtainable by providing improved airfoil cooling. Also, since the compressed air used to cool these airfoils is taken from the compressor, the use of a minimal amount of compressor bleed off air for the airfoil cooling will also increase the engine efficiency.
Airfoil cooling is also important in increasing the life of the airfoils. Hot spots can occur on sections of the airfoils that are not adequately cooled. These hot spots can cause oxidation that will lead to shortened life for the airfoil. Blade tips are especially subject to hot spots since it is nearly impossible to total eliminate the gap between the rotating blade tip and the stationary shroud that forms the gap. Without any gas, blade tip rubbing will occur which leads to other problems. Because of the presence of the tip gap, the hot gas can flow through the gap and expose the blade tip surface to the extreme high temperatures of the gas flow. Therefore, adequate blade tip cooling is also required to reduce hot gas flow leakage and to control metal temperature in order to increase part life.
Airfoils surfaces exposed to the high temperature gas flow are typically coated with a thermal barrier coating or TBC in order to allow for even higher temperatures. As the TBC technology improves, more industrial gas turbine (LOT) blades are applied with a thicker or low conductivity TBC. Cooling flow demand has been gradually reduced. FIG. 1 shows a prior art first stage turbine blade external pressure profile. As indicated by this figure, the forward region of the pressure side surface is exposed to high hot gas static pressure while the entire suction side of the airfoil is at much lower hot gas static pressure than the pressure side. As a result, there is not sufficient cooling flow for the design to split the total cooling flow into two or three flow circuits and utilize the forward flowing serpentine cooling design. Serpentine flow cooling circuits provide higher cooling capabilities than several straight channels in the airfoil because the overall cooling passage length is increased due to the looping of the circuit up and down the airfoil. Cooling flow for the blade leading and trailing edges has to be combined with the mid-chord flow circuit to form a single 5-pass serpentine flow circuit. However, for the forward 5-pass serpentine flow circuit with total blade cooling flow back flow margin (BFM) may become a design issue.
One prior art airfoil cooling design is the triple pass serpentine flow cooling design of FIG. 2 that includes a forward flowing 3-pass or triple pass circuit and an aft flowing triple pass serpentine circuit. The forward flowing serpentine flow circuit normally is designed in conjunction with the leading backside impingement plus a showerhead and pressure side and suction side film discharge cooling holes. The aft, flowing serpentine flow circuit is designed in conjunction with the airfoil trailing edge discharge cooling holes. This type of cooling flow circuit is called a dual triple pass cooling design. FIG. 3 shows a diagram view of the flow paths of the FIG. 2 circuit.
An alternative prior art cooling design utilizes the dual triple pass serpentine flow circuits for a high operating gas temperature is shown in FIG. 4 and is called the “Cold Bridge” cooling design. FIG. 5 shows a diagram view of this cooling circuit. In this particular cooling design, the leading edge airfoil is cooled with a self-contained flow circuit. The airfoil mid-chord section is cooled with a pair of triple pass serpentine flow circuits. However, both of the triple pass serpentine cooling flow circuits are flowing forward instead of one flowing forward and the other flowing aft-ward like in the “warm bridge” design of FIG. 2. For the first forward flowing triple pass serpentine cooling design used in the airfoil mid-chord region, the cooling air flows toward and discharges into the high hot gas side pressure section of the pressure side. In order to satisfy the back flow margin (BFM) criteria, a high cooling supply pressure is needed (to prevent ingestion of the hot gas flow into the airfoil interior through the film cooling holes) for this particular design, and thus inducing a high leakage flow. Since the last up-pass of the triple pass serpentine cavities provide film cooling air for both sides of the airfoil, in order to satisfy the back flow margin criteria for the pressure side film row the internal cavity pressure must be approximately 10% higher than the pressure side hot gas pressure which will result in over-pressuring the airfoil suction side film holes.
The second forward flowing serpentine flow circuit of FIG. 4 is designed in conjunction with the airfoil trailing edge discharge cooling holes. Cooling air for the airfoil trailing edge cooling is bled off from the triple pass serpentine first up pass cooling supply channel first to provide the airfoil trailing edge region cooling prior to any heating of the cooling air. In this particular cooling design, it achieves a direct trailing edge cooling with fresh cooling air and thus achieves the high temperature cooling design requirements.
The cooling circuit in FIG. 6 shows another prior art (1+3+3) serpentine flow cooling circuit for the first stage turbine blade. The flow path for the 1+3+3 flow circuit is shown in FIG. 7. For the second triple pass serpentine flow cooling circuit, the cooling air is used to provide cooling for the airfoil mid-chord region first, similar to the warm bridge design of FIG. 2. The cooling air then flows toward the airfoil trailing edge and discharges through the airfoil trailing edge cooling holes to provide cooling for the airfoil trailing edge corner.
For a low cooling flow designed and high temperature turbine blade that is coated with a TBC, a cooling design with cooling flow split three ways becomes unfeasible. Cooling air for the airfoil leading edge and trailing edge has to be combined into the serpentine flow circuit. However, cooling air flowing toward the airfoil leading edge with heated air will not be able to provide adequate leading edge region cooling. The forward flowing triple pass circuit for the airfoil forward region has to be designed as an aft flowing serpentine flow cooling circuit.